The present invention relates to a blade for a gas turbine. More specifically, the present invention relates to the cooling of a gas turbine blade shroud.
A gas turbine is typically comprised of a compressor section that produces compressed air. Fuel is then mixed with and burned in a portion of this compressed air in one or more combustors, thereby producing a hot compressed gas. The hot compressed gas is then expanded in a turbine section to produce rotating shaft power.
The turbine section typically employs a plurality of alternating rows of stationary vanes and rotating blades. Each of the rotating blades has an airfoil portion and a root portion by which it is affixed to a rotor.
Since the blades are exposed to the hot gas discharging from the combustors, the cooling of these components is of the utmost importance. Traditionally, cooling is accomplished by extracting a portion of the compressed air from the compressor, which may or may not then be cooled, and directing it to the turbine section, thereby bypassing the combustors. After introduction into the turbine, the cooling air flows through radial passages formed in the airfoil portions of the blades. Typically, the radial passages discharge the cooling air radially outward at the blade tip. In addition, a number of small passages may extend from one or more of the radial passages and direct the cooling air over the surfaces of the airfoils, such as the leading and trailing edges or the suction and pressure surfaces. After the cooling air exits the blade it enters and mixes with the hot gas flowing through the turbine section.
In some cases, turbine blades incorporate shrouds that project outwardly from the airfoil at the blade tip. Such shrouds serve to prevent hot gas leakage past the blade tips. In addition, if the shrouds are of the interlocking type, they may also serve to reduce blade vibration.
The approach to blade cooling discussed above provides adequate cooling to the airfoil portions of the blades. However, typically, no cooling air was specifically designated for use in cooling the blade shroud. Although the portion of the cooling air discharged from the radial passages at the blade tip flows over the radially outward facing surface of the shroud, so as to provide a measure of film cooling, experience has shown that this film cooling is insufficient to adequately cool the shroud. This is the result of the fact that by the time the cooling air exits the radial passages at the blade tip it has been heated to a temperature that may approach that of the hot gas flowing over the blade. As a result, creep and creep failures can occur in the blade shrouds due to operation at excessive temperatures.
One possible solution is to increase the amount of cooling air flowing through the radial passages and, therefore, avoid overheating of the cooling air by the time it reaches the blade tip. However, the increase in cooling air flow rate necessary to ensure a relatively low amount of heatup while flowing through the radial passages would be very large. Such a large increase in cooling air flow is undesirable. Although such cooling air enters the hot gas flowing through the turbine section when it exits at the blade tip, little useful work is obtained from the cooling air, since it is not subject to heat up in the combustion section. Thus, to achieve high efficiency, it is crucial that the use of cooling air be kept to a minimum.
It is therefore desirable to provide a scheme for cooling the shroud portion of the rotating blades in a gas turbine using a minimum of cooling air.